US4441313A - Mounting devices for multiple flow turbojet engines - Google Patents

Mounting devices for multiple flow turbojet engines Download PDF

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Publication number
US4441313A
US4441313A US06/204,582 US20458280A US4441313A US 4441313 A US4441313 A US 4441313A US 20458280 A US20458280 A US 20458280A US 4441313 A US4441313 A US 4441313A
Authority
US
United States
Prior art keywords
arm
shroud
exhaust casing
turbine
tierods
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US06/204,582
Other languages
English (en)
Inventor
Raymond J. M. Joubert
Jean G. Bouiller
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA filed Critical Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Assigned to SOCIETE NATIONALE D'ETUDE DE CONSTRUCTION DE MOTEURS D'AVIATION, "S.N.E.C.M.A." reassignment SOCIETE NATIONALE D'ETUDE DE CONSTRUCTION DE MOTEURS D'AVIATION, "S.N.E.C.M.A." ASSIGNMENT OF ASSIGNORS INTEREST. Assignors: BOUILLER JEAN G., JOUBERT RAYMON J. M.
Application granted granted Critical
Publication of US4441313A publication Critical patent/US4441313A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/40—Arrangements for mounting power plants in aircraft
    • B64D27/404—Suspension arrangements specially adapted for supporting vertical loads
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/20—Mounting or supporting of plant; Accommodating heat expansion or creep
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
    • F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
    • F05D2230/644—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins for adjusting the position or the alignment, e.g. wedges or eccenters
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T74/00—Machine element or mechanism
    • Y10T74/18—Mechanical movements
    • Y10T74/18888—Reciprocating to or from oscillating
    • Y10T74/1892—Lever and slide
    • Y10T74/18952—Lever and slide toggle transmissions
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T74/00—Machine element or mechanism
    • Y10T74/20—Control lever and linkage systems
    • Y10T74/20576—Elements
    • Y10T74/20582—Levers
    • Y10T74/20588—Levers toggle

Definitions

  • the invention concerns aircraft turbojet engines and more particularly relates to a mounting device, on the aircraft, having a turbine, said turbine having an axis of the rear part of a multiple flow turbojet engine having at least two concentric flows.
  • the rear mount is applied to the exhaust casing of the turbine, which constitutes the rear part of the stator shroud and supports the rear bearing.
  • the turbojet engine is of the double of triple flow type, the exhaust casing--which contains the central primary flow--is suspended in the shroud of the secondary flow, which itself is mounted directly on the supporting structure.
  • This suspension of the exhaust casing is effected by means of a plurality connecting elements distributed around the casing. Each of these elements consists of an arm, with a cross section profiled to accommodate the flow.
  • Each connecting element is equipped with first fastening means for connection to the shroud of the secondary flow and with second fastening means for connecting to the exhaust casing.
  • the secondary flow shroud is rigidly joined to the supporting structure
  • the exhaust casing is rigidly joined to the shroud of the secondary flow
  • the device of the invention intends to secure the exhaust casing of a turbojet engine having at least two flows to a rigid support structure joined to an aircraft.
  • the two flows may be, for example, a central primary flow contained by the exhaust casing and a secondary flow contained by a shroud surrounding the exhaust casing and attached to the support structure.
  • the device is of a type comprising a plurality of connecting elements distributed around the exhaust casing and each comprising an arm oriented in a plane transverse to the axis of the rotor of the turbine, first fastening means to attach the arm to the shroud of the secondary flow casing, and second fastening means to attach the arm to the exhaust casing, said device being essentially characterized in that the first fastening means of each arm comprise on the one hand an articulation of the arm joining one section of the arm with the shroud of the secondary flow while leaving the arm free to rotate in said transverse plane, and on the other hand, a tie of variable length oriented in said transverse plane and joined respectively to the secondary flow shroud and to another section of the arm by means of two tierod articulations free to rotate in said transverse plane.
  • each tierod is provided with means to render it elastically deformable. Later, it will be seen how this disposition eases the consequences of certain grave failures of the turbojet engine.
  • a further advantage of the invention results from the second joining means of each arm conferring on the arm, relative to the exhaust casing, a freedom of rotation in said transverse plane and a freedom of translation in the direction of said plane passing in the vicinity of the turbine axis. It will be seen later herein how this disposition eases in particular the consequences of the differential expansion appearing in the course of operation.
  • FIG. 1 is a transverse section of a triple flow turbojet engine equipped with the device of the present invention
  • FIG. 2 is a partial axial section taken on the line 6--6 of FIG. 1;
  • FIG. 3 is a view on a larger scale of the suspension means shown in FIG. 1;
  • FIG. 4 is a view on a larger scale of one of the fastening elements shown in FIG. 1;
  • FIG. 5 is a top view of the fastening element shown in FIG. 4.
  • FIGS. 1 and 2 show a rigid, external ring 50 which constitutes the rear part of the shroud for the tertiary flow, a rigid intermediate ring 60 which constitutes the rear part of the shroud for the secondary flow and an internal ring formed by the exhaust casing 70 and surrounding the cone 71 of the rear bearing support.
  • the latter is secured to the ring 70 by means of a plurality of profiled arms 72.
  • the ring 50 and the ring 60 are linked rigidly, on the one hand by the rear frame 80 suspended by the tierods 81 from a case 82 (only its forward flange shown) integral with an aircraft strut, not shown, and on the other by a plurality of radial attach spacers 61.
  • the external ring 50 is joined to the forward body of the shroud of the tertiary flow by a tight, deformable joint consisting of an elastic bellows 62.
  • the tierods 81 and the case 82 are visible in FIG. 1 and not in FIG. 2.
  • the ring 70 is suspended in the ring 60 by means of four connecting elements 90 which shall be described later with reference to FIGS. 4 and 5.
  • the lower ends of the tierods 81 are connected with the frame 80 by axles 83, while the upper ends are connected to the case 82 by means of axles 84. If the system would be infinitely rigid, the two tierods, the flange of the case 82 and the frame 80 would form a nondeformable quadrangle because the rotation of the nondeformable assembly consisting of the rings 50 and 60 around its axis is made impossible by the forward link. However, the elastic deformations of the structures connected by the tierods 81 permit slight displacements inside the passages 85 of the tierods 81 provided in the frame 80. Each body 81 of the tierods is surrounded by a cylindrical bellows 86 which connects the body with the outlet of the corresponding passage in order to ensure the tightness of the tertiary jet flow.
  • FIGS. 4 and 5 represent schematically one of the connecting elements 90 shown in FIG. 2.
  • the arm 91 with its profiled box extends through a passage 63 in the ring 60. It has an extension 92 by which it is pivoted on an axle 64 located in the passage 63 with its axis oriented parallel to the axis of the turbine. The arm 91 is thus free to rotate in a plane transverse to the axis of the turbine.
  • the arm 91 carries at its external end (i.e., the end in the tertiary flow path) a plurality of ears 93 traversed by an axle 94 parallel to the axle 64.
  • the stationary attach spacer 61 of the shape shown, has in its medium section, the ears 65 which extend toward the ears 93 and which are traversed by an axle 66 that is parallel to the axle 64.
  • axle 94 supports by their ends a first pair of tierods 95.
  • axle 66 supports by their ends a second pair of tierods 95.
  • a pair of axles 96 connects a tierod of one pair of tierods with a tierod of the other pair.
  • the axles 66 and 94 are connected with each other by a plurality of identical, deformable parallelograms, each consisting of four tierods 95.
  • Spaces 97, placed on the axles 96 between the pairs of tierods ensure the bracing of these parallelograms, the assembly whereof constitutes a deformable, dihedral parallelogram providing the connection between an orientable arm 91 and a stationary attach spacer 61.
  • the ends of each axle 96 extend outwardly of both sides of the tierods 95 forming the bases of said parallelogram and each end bears a cheek member 98.
  • the two cheek members which flank the same base are joined by two bolts 99 parallel to each other, placed on either side of the two axles 96 and entering threaded holes provided in said cheek members. At opposite ends of the bolts 99, the threading is inversed (of reverse hand), together with the corresponding threading of the cheek members 98, so that the inclination of the arms 91 may be adjusted during assembly by rotating the bolts 99.
  • each arm 91 (placed in the secondary flow path) has a cylindrical stud 101 (shown by a dotted line in FIG. 5), which engages with a slight friction the reamed hole provided in a swivel 102.
  • the latter is inserted in a spherical housing provided in a bearing 103, which itself is capable of sliding in a channel 73, integral with the shroud 70 and oriented parallel to the axis of the turbine.
  • the different elements located in a plane transverse to the axis of the turbine and which have been described hereinabove with reference to FIGS. 4 and 5, thus cooperate to regulate and impart to each arm 91 a given orientation in a plane parallel to the axis of the turbine and to provide said arm, relative to the exhaust casing 70, with a freedom of position in said plane, a freedom of radial translation (i.e., in a direction passing through the axis of the turbine) and a freedom of axial translation (i.e., in a direction parallel to said axis).
  • each arm 91 may be force mounted in the swivel 102 and the freedom of axial translation of the arm 91 relative to the shroud 70 is assured by means of the sliding of the swivel bearing 103 which would be made capable of sliding in both the radial and axial direction in the channel 73.
  • Suitable values of stiffness and inertia of the connection effected by the tierods 95 are obtained by assigning to said tierods definite dimensions in order to ensure the absorption of vertical and horizontal stresses, together with the force couples in this plane. It may be noted that the different parallelogram also represent variable braces to adjust the connection between the two shrouds constituting the stator.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Control Of Turbines (AREA)
  • Exhaust Gas After Treatment (AREA)
  • Physical Or Chemical Processes And Apparatus (AREA)
US06/204,582 1979-11-12 1980-11-06 Mounting devices for multiple flow turbojet engines Expired - Lifetime US4441313A (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR7927771 1979-11-12
FR7927771A FR2469566A1 (fr) 1979-11-12 1979-11-12 Perfectionnements aux dispositifs de fixation de turboreacteurs multiflux

Publications (1)

Publication Number Publication Date
US4441313A true US4441313A (en) 1984-04-10

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ID=9231536

Family Applications (1)

Application Number Title Priority Date Filing Date
US06/204,582 Expired - Lifetime US4441313A (en) 1979-11-12 1980-11-06 Mounting devices for multiple flow turbojet engines

Country Status (4)

Country Link
US (1) US4441313A (cg-RX-API-DMAC10.html)
EP (1) EP0028970B1 (cg-RX-API-DMAC10.html)
DE (1) DE3067468D1 (cg-RX-API-DMAC10.html)
FR (1) FR2469566A1 (cg-RX-API-DMAC10.html)

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4571936A (en) * 1985-07-10 1986-02-25 The United States Of America As Represented By The Secretary Of The Air Force Length adjustable strut link with low aerodynamic drag
US4658579A (en) * 1983-07-14 1987-04-21 United Technologies Corporation Load sharing for engine nacelle
US4875655A (en) * 1987-03-13 1989-10-24 General Electric Company Vibration isolating engine mount
US5028001A (en) * 1987-03-13 1991-07-02 General Electric Company Method of vibration isolating an aircraft engine
US5351930A (en) * 1992-08-11 1994-10-04 Lord Corporation Mounting for engines and the like
US5437412A (en) * 1993-06-23 1995-08-01 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) Variable geometry jet engine exhaust nozzle
WO2002012680A1 (de) * 2000-08-03 2002-02-14 Mtu Aero Engines Gmbh Aufhängung
US20070031247A1 (en) * 2005-08-05 2007-02-08 Siemens Westinghouse Power Corporation Radially expanding turbine engine exhaust cylinder interface
US20080223992A1 (en) * 2007-03-16 2008-09-18 The Boeing Company High-Performance Low-Noise Aircraft Exhaust Systems and Methods
US7581922B1 (en) * 2005-05-16 2009-09-01 Mitsubishi Heavy Industries, Ltd. Turbine casing structure
US20090229244A1 (en) * 2006-05-04 2009-09-17 Volvo Aero Corporation Device for pivoting at least one pivotable element in a gas turbine engine
US20100229569A1 (en) * 2006-06-09 2010-09-16 Charles Hollimon Engine Exhaust System
US20110131948A1 (en) * 2006-06-09 2011-06-09 Charles Hollimon Engine Exhaust System with Directional Nozzle
US8733693B2 (en) 2008-02-28 2014-05-27 Airbus Operations S.A.S. Aircraft engine assembly comprising an annular load-transfer structure surrounding the central casing of a turbojet engine
CN104948301A (zh) * 2015-05-20 2015-09-30 西安交通大学 一种燃气轮机的热端支承
US20160003192A1 (en) * 2013-03-06 2016-01-07 United Technologies Coproration Exhaust system having a flow path liner supported by structural duct segments
US9366186B2 (en) 2010-07-30 2016-06-14 Snecma Lateral turbojet improved in order to limit the deformation thereof
CN107585318A (zh) * 2016-07-08 2018-01-16 空中客车运营简化股份公司 飞行器发动机组件
US10518891B2 (en) * 2014-11-21 2019-12-31 General Electric Company Turbine engine assembly and method of manufacturing thereof
US20220316358A1 (en) * 2021-04-02 2022-10-06 Mitsubishi Heavy Industries, Ltd. Gas turbine
US11788436B2 (en) * 2019-04-16 2023-10-17 Safran Aircraft Engines Exhaust casing of a turbomachine with improved aerodynamics

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH0717238B2 (ja) * 1983-07-14 1995-03-01 ユナイテツド・テクノロジ−ズ・コ−ポレイシヨン 荷重分担アダプタを備えたガスタービンエンジン
GB2192233B (en) * 1986-07-02 1990-11-28 Rolls Royce Plc A gas turbine engine load transfer structure
US7814753B2 (en) * 2006-07-25 2010-10-19 United Technologies Corporation Low profile attachment hanger system for a cooling liner within a gas turbine engine swivel exhaust duct
EP1887209A3 (en) * 2006-07-25 2010-11-17 United Technologies Corporation Hanger system for a cooling liner within a gas turbine engine exhaust duct
FR2905975B1 (fr) * 2006-09-20 2008-12-05 Snecma Sa Conduite de soufflante pour une turbomachine.
FR3060532B1 (fr) * 2016-12-20 2019-05-31 Airbus Operations Ensemble moteur pour aeronef, comprenant des dispositifs elastiques souples de transmission d'efforts entre des capots de nacelle et une structure annulaire reliee au carter moteur par des bielles

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US2936978A (en) * 1957-03-29 1960-05-17 United Aircraft Corp Rear engine mount
US2936999A (en) * 1956-12-07 1960-05-17 United Aircraft Corp Tangential bearing supports
US3412560A (en) * 1966-08-03 1968-11-26 Gen Motors Corp Jet propulsion engine with cooled combustion chamber, fuel heater, and induced air-flow
FR2102187A1 (cg-RX-API-DMAC10.html) * 1970-08-11 1972-04-07 Secretaire Etat Royaume
US3886737A (en) * 1972-08-22 1975-06-03 Mtu Muenchen Gmbh Turbojet engines of multi-shaft and multi-flow construction

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DE961151C (de) * 1956-02-18 1957-04-04 W Blume Ingenieurbuero Fuer Le Anordnung von TL-Triebwerken an Flugzeugen
FR1198215A (fr) * 1958-06-10 1959-12-04 M Le Ministre De La Defense Na Dispositif d'accouplement réglable, notamment de contre-fiches arrière de turbo-moteurs à gaz
US3327965A (en) * 1965-09-27 1967-06-27 Douglas Aircraft Inc Flexible engine pylon

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Publication number Priority date Publication date Assignee Title
US2936999A (en) * 1956-12-07 1960-05-17 United Aircraft Corp Tangential bearing supports
US2936978A (en) * 1957-03-29 1960-05-17 United Aircraft Corp Rear engine mount
US3412560A (en) * 1966-08-03 1968-11-26 Gen Motors Corp Jet propulsion engine with cooled combustion chamber, fuel heater, and induced air-flow
FR2102187A1 (cg-RX-API-DMAC10.html) * 1970-08-11 1972-04-07 Secretaire Etat Royaume
US3886737A (en) * 1972-08-22 1975-06-03 Mtu Muenchen Gmbh Turbojet engines of multi-shaft and multi-flow construction

Cited By (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4658579A (en) * 1983-07-14 1987-04-21 United Technologies Corporation Load sharing for engine nacelle
US4571936A (en) * 1985-07-10 1986-02-25 The United States Of America As Represented By The Secretary Of The Air Force Length adjustable strut link with low aerodynamic drag
US4875655A (en) * 1987-03-13 1989-10-24 General Electric Company Vibration isolating engine mount
US5028001A (en) * 1987-03-13 1991-07-02 General Electric Company Method of vibration isolating an aircraft engine
US5351930A (en) * 1992-08-11 1994-10-04 Lord Corporation Mounting for engines and the like
US5437412A (en) * 1993-06-23 1995-08-01 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) Variable geometry jet engine exhaust nozzle
WO2002012680A1 (de) * 2000-08-03 2002-02-14 Mtu Aero Engines Gmbh Aufhängung
US6752591B2 (en) 2000-08-03 2004-06-22 Mtu Aero Engines Gmbh Suspension
US7581922B1 (en) * 2005-05-16 2009-09-01 Mitsubishi Heavy Industries, Ltd. Turbine casing structure
US20090226313A1 (en) * 2005-05-16 2009-09-10 Mitsubishi Heavy Industries, Ltd. Turbine casing structure
US20070031247A1 (en) * 2005-08-05 2007-02-08 Siemens Westinghouse Power Corporation Radially expanding turbine engine exhaust cylinder interface
US7584621B2 (en) * 2005-08-05 2009-09-08 Siemens Energy, Inc. Radially expanding turbine engine exhaust cylinder interface
US8347601B2 (en) * 2006-05-04 2013-01-08 Volvo Aero Corporation Device for pivoting at least one pivotable element in a gas turbine engine
US20090229244A1 (en) * 2006-05-04 2009-09-17 Volvo Aero Corporation Device for pivoting at least one pivotable element in a gas turbine engine
US20110131948A1 (en) * 2006-06-09 2011-06-09 Charles Hollimon Engine Exhaust System with Directional Nozzle
US20100229569A1 (en) * 2006-06-09 2010-09-16 Charles Hollimon Engine Exhaust System
US8123159B2 (en) * 2006-06-09 2012-02-28 Textron Innovations Inc. Engine exhaust system
US8132755B2 (en) * 2006-06-09 2012-03-13 Bell Helicopter Textron Inc. Engine exhaust system with directional nozzle
US7726609B2 (en) 2007-03-16 2010-06-01 The Boeing Company High-performance low-noise aircraft exhaust systems and methods
US20080223992A1 (en) * 2007-03-16 2008-09-18 The Boeing Company High-Performance Low-Noise Aircraft Exhaust Systems and Methods
US8733693B2 (en) 2008-02-28 2014-05-27 Airbus Operations S.A.S. Aircraft engine assembly comprising an annular load-transfer structure surrounding the central casing of a turbojet engine
US9366186B2 (en) 2010-07-30 2016-06-14 Snecma Lateral turbojet improved in order to limit the deformation thereof
US20160003192A1 (en) * 2013-03-06 2016-01-07 United Technologies Coproration Exhaust system having a flow path liner supported by structural duct segments
US9982628B2 (en) * 2013-03-06 2018-05-29 United Technologies Corporation Exhaust system having a flow path liner supported by structural duct segments
US10518891B2 (en) * 2014-11-21 2019-12-31 General Electric Company Turbine engine assembly and method of manufacturing thereof
US11485508B2 (en) 2014-11-21 2022-11-01 General Electric Company Turbine engine assembly and method of manufacturing thereof
CN104948301A (zh) * 2015-05-20 2015-09-30 西安交通大学 一种燃气轮机的热端支承
CN107585318A (zh) * 2016-07-08 2018-01-16 空中客车运营简化股份公司 飞行器发动机组件
CN107585318B (zh) * 2016-07-08 2022-04-08 空中客车运营简化股份公司 飞行器发动机组件
US11788436B2 (en) * 2019-04-16 2023-10-17 Safran Aircraft Engines Exhaust casing of a turbomachine with improved aerodynamics
US20220316358A1 (en) * 2021-04-02 2022-10-06 Mitsubishi Heavy Industries, Ltd. Gas turbine
US11708771B2 (en) * 2021-04-02 2023-07-25 Mitsubishi Heavy Industries, Ltd. Gas turbine with reduced axial displacement under thermal expansion

Also Published As

Publication number Publication date
FR2469566A1 (fr) 1981-05-22
EP0028970A1 (fr) 1981-05-20
FR2469566B1 (cg-RX-API-DMAC10.html) 1984-11-09
EP0028970B1 (fr) 1984-04-11
DE3067468D1 (en) 1984-05-17

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AS Assignment

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